The term “rotating blade assembly” when used in the specification is to be understood as referring to an assembly including a rotor mounting a plurality of blades, some of which blades are axially spaced relative to each other on the rotor. The rotor is rotatably mounted in a casing, with a number of stators provided on the casing, with the blades passing between the stators during rotation. Such assemblies may typically be used as turbines or compressors, and particularly in gas turbine engines such as are used on aircraft.
Referring to FIG. 1, a gas turbine engine is generally indicated at 10 and comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, a combustor 15, a turbine arrangement comprising a high pressure turbine 16, an intermediate pressure turbine 17 and a low pressure turbine 18, and an exhaust nozzle 19.
The gas turbine engine 10 operates in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 which produce two air flows: a first air flow into the intermediate pressure compressor 13 and a second air flow which provides propulsive thrust. The intermediate pressure compressor compresses the air flow directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high pressure compressor 14 is directed into the combustor 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive, the high, intermediate and low pressure turbines 16, 17 and 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low pressure turbines 16, 17 and 18 respectively drive the high and intermediate pressure compressors 14 and 13 and the fan 12 by suitable interconnecting shafts 26, 28, and 30.
In view of the above it will be understood that the provision of rotating blade assemblies for gas turbine engines requires an intricate and relatively complex assembly of blades on a rotor disk or core. A traditional blade assembly such as that used in a low pressure turbine module comprises a number of individual turbine disks which are bolted together on final assembly to produce the low pressure turbine module. The individual blades are secured to the disks and then each disk is bolted to adjacent disks by a bolted flange joint provided on a drive element. Each disk will require individual balancing.
In order to achieve the above assembly, each turbine disk requires machining after a forging process. The forging process to allow machining must include the drive element which means that the forged section is generally much larger in cross section than the final machined part. This means there is a significant additional volume of material used, and an additional machining process which adds to cost. Clearly, reducing the amount of material utilised would provide significant benefits with regard to raw material costs, particularly as most materials used will be relatively expensive in order to meet the temperature and other requirements within the blade assembly.
It is not uncommon for the drive element along with corresponding flanges, nuts, bolts and washers for a blade assembly to weigh as much as 20 kg.